The airframe in an aircraft comprises the main structural body designing the shape and the structural behaviour of said aircraft. At present, composite materials are widely used in aircraft design, thus also being used for forming the frames of aircraft.
Looking at an aircraft fuselage, a typical structure includes skin and frames, each of them with its different function and performance. An interface frame for instance is located between two sections. It is typically very demanded with regard to stresses and structural behaviour, thus having to withstand very high loads. For this reason, the interface frame has typically been made in a metallic material, in order to obtain good mechanical resistance and appropriate tolerances. The fabrication of these interface frames of metallic material is both costly and time consuming, also requiring a high number of pieces effecting the joints of the different metallic segments, fabricated separately. Besides, corrosion plays a fundamental role when metallic materials are used.
Moreover, a metallic material is an isotropic material, so the design of metallic interface frames is made in such a way that the same kind of material having the same resistance is used all over the interface frame. Therefore, more material is needed than in a case in which an anisotropic material is used, such as a composite material: in such a case, the material is distributed so as to provide higher resistance only where it is needed.
Lastly, now that composites materials are more and more used, specially in skins, hybrid structures containing composites and metallic materials show many integration problems: thermal expansions differences, galvanic corrosion, disparity in tolerance concepts, analysis procedures, etc.
To make best use of composite materials, an all-composite structure shall be aimed.
Document US 2009277994 discloses a hybrid airframe, comprising structural components made in a metallic material and non-structural components made in composite material. The airframe also comprises metallic frames, reinforced by means of longitudinal composite stringers, metallic joint members and metallic ties. The main disadvantage of such structure is that the weight is higher than in a case in which only composite material is used. Moreover, the manufacturing process and joints of these elements is costly and time consuming.
Document EP 1030807 describes a composites structural solution for a rear pressure frame for an aircraft. Pressures frames are very special structural parts, as they are the interface between pressurized and unpressurized fuselage sections and thus they must withstand very specific pressure loads. However, this solution cannot be applied to other structural frames, where all this composite closed surface is not needed and means too much of unnecessary weight.
Document WO 2009/129007 discloses a method for manufacturing composite material frames for aircraft having multiple legs (webs). This manufacturing method covers on more way to achieve the traditional structural concept of a frame, but it does not present any innovative structural concept.
Document GB 2268461 discloses a hybrid frame for an aircraft, comprising composite external and internal covers, together with reinforcing elements made in a metallic material. This document gives a manufacturing solution to a central fuselage and flying surfaces attached in an integrated way. The joints of these two components are mainly the object of the GB 2268461 invention.
It would thus be desirable to provide a structural solution for a frame of the fuselage of an aircraft, fully made in composite material and without the need of riveting different pieces to provide the whole structure of the frame, thus being provided a more effective fabrication method of said interface frame.
The present invention is oriented towards this need.